Additively manufactured core for use in casting an internal cooling circuit of a gas turbine engine component

ABSTRACT

A core for use in casting an internal cooling circuit within a gas turbine engine component includes a base core portion and an additive core portion additively manufactured to the base core portion. A method of manufacturing a core for use in casting an internal cooling circuit within a gas turbine engine component including additively manufacturing an additive core portion to a base core portion.

BACKGROUND

The present disclosure relates to additive manufacturing and, moreparticularly, to a core with an additively manufactured portion for usein casting an internal cooling circuit within a gas turbine enginecomponent.

Gas turbine engines typically include a compressor section to pressurizeairflow, a combustor section to burn a hydrocarbon fuel in the presenceof the pressurized air, and a turbine section to extract energy from theresultant combustion gases. Gas path components, such as turbine blades,often include airfoil cooling that may be accomplished by external filmcooling, internal air impingement and forced convection eitherseparately or in combination.

Advances in casting facilitate significantly smaller and more complexpassages to accommodate elevated temperatures with a reduced flow ofcooling air, yet relatively small features remain difficult, if notimpossible, to cast via a conventional core casting process in anefficient and repeatable manner.

SUMMARY

A core for use in casting an internal cooling circuit within a gasturbine engine component according to one disclosed non-limitingembodiment of the present disclosure can include a base core portion;and an additive core portion additively manufactured to the base coreportion.

A further embodiment of the present disclosure may include, wherein thebase core portion is manufactured of a first material and the additivecore portion is manufactured of a second material, the first materialdifferent than the second material.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the base core portion is manufactured of a ceramicmaterial.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the base core portion is cast.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the additive core portion is additivelymanufactured of molybdenum.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the additive core portion forms a multiple oftrailing edge crossovers.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the additive core portion is additivelymanufactured onto the base core portion.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the additive core portion includes a multiple ofpin-shaped protrusions.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein each of the multiple of pin-shaped protrusions areadditively manufactured into a respective recess formed in the base coreportion.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the base core portion is more crushable than theadditive core portion.

A method of manufacturing a core for use in casting an internal coolingcircuit within a gas turbine engine component, the method according toanother disclosed non-limiting embodiment of the present disclosure caninclude additively manufacturing an additive core portion to a base coreportion.

A further embodiment of any of the embodiments of the present disclosuremay include casting the base core portion.

A further embodiment of any of the embodiments of the present disclosuremay include casting the base core portion of a first material andadditively manufacturing the additive core portion of a second material,the first material different than the second material.

A further embodiment of any of the embodiments of the present disclosuremay include forming the additive core portion to form a multiple oftrailing edge crossovers.

A further embodiment of any of the embodiments of the present disclosuremay include additively manufacturing the additive core portion to form amultiple of pin-shaped protrusions.

A further embodiment of any of the embodiments of the present disclosuremay include forming the additive core portion onto the base coreportion.

A further embodiment of any of the embodiments of the present disclosuremay include firing the additive core portion and the base core portion.

A further embodiment of any of the embodiments of the present disclosuremay include additively manufacturing the additive core portion to form atrailing edge portion of the core.

A further embodiment of any of the embodiments of the present disclosuremay include positioning the core within a shell.

A further embodiment of any of the embodiments of the present disclosuremay include, wherein the additive core portion at least partiallycontacts the shell.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture;

FIG. 2 is an enlarged schematic cross-section of an engine turbinesection;

FIG. 3 is a perspective view of an airfoil as an example component;

FIG. 4 is a schematic cross-section view of the airfoil of FIG. 4showing the internal architecture;

FIG. 5 is a schematic partial fragmentary view of a mold with ceramiccore that is additively manufactured according to another disclosedprocess for casting of an airfoil;

FIG. 6 is a method to at least partially additively manufacture a coreaccording to one disclosed non-limiting embodiment;

FIG. 7 is a perspective view of a portion of a core;

FIG. 8 is a perspective view of a core with an additively manufacturedportion;

FIG. 9 is an expanded sectional view of an additively manufactured corefeature features according to another disclosed non-limiting embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flowpath while the compressor section 24 drives airalong a core flowpath for compression and communication into thecombustor section 26 then expansion through the turbine section 28.Although depicted as a turbofan in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbine engine architectures such as turbojets,turboshafts, and three-spool (plus fan) turbofans.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Xrelative to an engine static structure 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor (“LPC”) 44 and a lowpressure turbine (“LPT”) 46. The inner shaft 40 drives the fan 42directly or through a geared architecture 48 to drive the fan 42 at alower speed than the low spool 30. An exemplary reduction transmissionis an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54. Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis X whichis collinear with their longitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with thefuel and burned in the combustor 56, then expanded over the HPT 54 andthe LPT 46. The turbines 54, 46 rotationally drive the respective lowspool 30 and high spool 32 in response to the expansion. The main engineshafts 40, 50 are supported at a plurality of points by bearingstructures 38 within the static structure 36. It should be understoodthat various bearing structures 38 at various locations mayalternatively or additionally be provided.

With reference to FIG. 2, an enlarged schematic view of a portion of theturbine section 28 is shown by way of example; however, other enginesections will also benefit herefrom. A full ring shroud assembly 60within the engine case structure 36 supports a blade outer air seal(BOAS) assembly 62 with a multiple of circumferentially distributed BOAS64 proximate to a rotor assembly 66 (one schematically shown).

The full ring shroud assembly 60 and the BOAS assembly 62 are axiallydisposed between a forward stationary vane ring 68 and an aft stationaryvane ring 70. Each vane ring 68, 70 includes an array of vanes 72, 74that extend between a respective inner vane platform 76, 78 and an outervane platform 80, 82. The outer vane platforms 80, 82 are attached tothe engine case structure 36.

The rotor assembly 66 includes an array of blades 84 circumferentiallydisposed around a disk 86. Each blade 84 includes a root 88, a platform90 and an airfoil 92 (also shown in FIG. 3). The blade roots 88 arereceived within a rim 94 of the disk 86 and the airfoils 92 extendradially outward such that a tip 96 of each airfoil 92 is closest to theblade outer air seal (BOAS) assembly 62. The platform 90 separates a gaspath side inclusive of the airfoil 92 and a non-gas path side inclusiveof the root 88.

With reference to FIG. 3, the platform 90 generally separates the root88 and the airfoil 92 to define an inner boundary of a gas path. Theairfoil 92 defines a blade chord between a leading edge 98, which mayinclude various forward and/or aft sweep configurations, and a trailingedge 100. A first sidewall 102 that may be convex to define a suctionside, and a second sidewall 104 that may be concave to define a pressureside are joined at the leading edge 98 and at the axially spacedtrailing edge 100. The tip 96 extends between the sidewalls 102, 104opposite the platform 90. It should be appreciated that the tip 96 mayinclude a recessed portion.

To resist the high temperature stress environment in the gas path of aturbine engine, each blade 84 may be formed by casting. It should beappreciated that although a blade 84 with an internal cooling circuit110 (shown schematically; FIG. 4) will be described and illustrated indetail, other hot section components including, but not limited to,vanes, turbine shrouds, end walls, and other such components will alsobenefit here from.

With reference to FIG. 4, the internal cooling circuit 110 may include afeed passage 112 that communicates airflow into a trailing edge cavity114 within the airfoil 84. It should be appreciated that the internalcooling circuit 110 may be of various geometries, and include variousfeatures. The feed passage 112 in this embodiment is the aft mostpassage that communicates cooling air to the trailing edge cavity 114.The feed passage 112 generally receives cooling flow through at leastone inlet 116 within the base 118 of the root 88. It should beappreciated that various feed architecture; cavities, and passagewayarrangements will benefit herefrom.

The tip 96 and the trailing edge 100 bound the trailing edge cavity 114between the sidewalls 102, 104. The trailing edge cavity 114 includes amultiple of features 120. The features 120 in this disclosednon-limiting embodiment may include a multiple of pedestals 122, amultiple of strips 124, and a multiple of edge features 126. It shouldbe appreciated that although particular features are delineated withincertain general areas, the features may be otherwise arranged orintermingled and still not depart from the disclosure herein.

The pedestals 122 may be staggered and be of one or more shapes such ascircular, rectilinear, diamond and others. The pedestals 122 generateturbulence in the cooling airflow and hence advantageously increasesheat pick-up. The strip features 124 form a multiple of respectivetrailing edge crossovers 160. The trailing edge crossovers 160 extend tothe trailing edge 100. The edge features 126 define trailing edge exits162 through the trailing edge 100 such that the trailing edge 100 may beessentially discontinuous.

Generally, to form the internal cooling circuit 110, a core 200 ispositioned within a shell 202 (FIG. 5). The shell 202 defines the outersurface of the blade 84 while the core 200 forms the internal surfacessuch as that which defines the internal cooling circuit 110 (FIG. 4).That is, during the casting process, the core 200 fills a selectedvolume within the shell 202 that, when removed from the finished bladecasting, defines the internal cooling circuit 110 utilized for coolingairflow.

The shell 202 and the core 200 together define a mold 204 to cast thecomplex exterior and interior geometries that may be formed ofrefractory metals, ceramic, or hybrids thereof. The mold 204 operates asa melting unit and/or a die for a desired material that forms the blade84. The desired material may include but not be limited to a super alloyor other material such as nickel based super alloy, cobalt based superalloy, iron based super alloy, and mixtures thereof that is melted; amolten super alloy that is then solidified; or other material. Inanother non-limiting embodiment, the crucible may be directly filledwith a molten super alloy.

Alternatively, or in addition, a single crystal starter seed or grainselector may be utilized to enable a single crystal to form whensolidifying the component. The solidification may utilize a chill blockin a directional solidification furnace. The directional solidificationfurnace has a hot zone that may be induction heated and a cold zoneseparated by an isolation valve. The chill block and may be elevatedinto the hot zone and filled with molten super alloy. After the pour, orbeing molten, the chill plate may descend into the cold chamber causinga solid/liquid interface to advance from the partially molten starterseed in the form of a single crystallographic oriented component whoseorientation is dictated by the orientation of the starter seed. Castingis typically performed under an inert atmosphere or vacuum to preservethe purity of the casting.

Following solidification, the shell 202 may be broken away and the core200 may be removed from the solidified component by, for example,caustic leaching, to leave the finished single crystal component. Afterremoval, machining, surface treating, coating, or any other desirablefinishing operation may be performed to further finish the component.

With reference to FIG. 6, one disclosed non-limiting embodiment of amethod 300 to manufacture the core 200 initially includes manufacturinga base core portion 400 such as via injected or transfer molding withoptional finishing steps (step 310; FIG. 7). That is, the base coreportion 400 may be a semi-finished conventionally manufactured ceramiccore that is first manufactured. As one or more features are to beadditively manufactured, the base core portion 400 may be relativelyless complicated and need only be utilized to, for example, produce therelatively more coarse features of the internal cooling circuit 110.

Next, the base core portion 400 is fixtured into a bed of an additivemanufacturing machine (step 320). The shape of the base core portion 400is accounted for in the programming of the additive manufacturingmachine. That is, the programming is utilized to facilitate usage of thebase core portion 400 as a reference for the additive manufacturingmachine. It should be appreciated that in some embodiment, additivemanufacturing may be performed on one location of the base core portion400, then the base core portion 400 is fixtured in another orientationto perform additive manufacturing on another location of the base coreportion 400.

Next, an additive material fills the bed, and ceramic printing commencesupon or adjacent to the base core portion 400 to form an additive coreportion 402 that is additively manufactured to the base core portion 400(step 330; FIG. 8).

The additive core portion 402 may be readily manufactured with anadditive manufacturing process that includes but are not limited to,Sterolithography (SLA), Direct Selective Laser Sintering (DSLS),Electron Beam Sintering (EBS), Electron Beam Melting (EBM), LaserEngineered Net Shaping (LENS), Laser Net Shape Manufacturing (LNSM),Direct Metal Deposition (DMD), Laser Powder Bed Fusion (LPBF) andothers. Although particular additive manufacturing processes aredisclosed, those skilled in the art of manufacturing will recognize thatany other suitable rapid manufacturing methods using layer-by-layerconstruction or additive fabrication can alternatively be used.

The additive material for the additive core portion 402 may be amaterial different than that of the base core portion 400. For example,the base core portion 400 may be manufactured of silica, refractorymetal, alumina ceramic, or other such material, whereas the additivematerial may include molybdenum or other relatively more resilientmaterial. That is, the base core portion 400 is manufactured of arelatively crushable material, while the additive core portion 402 maybe manufactured of a relatively less crushable material to facilitatedefinition of the relatively fine portions of the internal coolingcircuit 110. For instance, the base core portion 400 may be producedwithout trailing edge features, which are later completed as theadditive core portion 402. It should be appreciated that various and/ormultiple additive core portions 402 may be formed.

Finally, the base core portion 400 and the additive core portion 402 arethen fired to complete the core 200 (step 340). That is, the base coreportion 400 and the additive core portion 402 are prepared for finaldisposition of the core 200 within the shell 202

With reference to FIG. 9, in another disclosed non-limiting embodiment,the additive core portion 402A may include a multiple of pin-shapedprotrusions 500 that are additively manufactured into a respectiverecess 502 formed in the base core portion 400A (one shown). Theprotrusions 500 are operable to form cooling apertures that extend intothough a bump on an internal surface of the airfoil wall to, forexample, provide an internal feature 120A to facilitate heat transfer.It should be appreciated that such high fidelity features may not beotherwise manufactured with conventional cast cores.

The additive core portion 402 readily facilities manufacture ofrelatively small features that are difficult if not impossible toproduce via a conventional core casting process in an efficientrepeatable manner.

The use of the terms “a,” “an,” “the,” and similar references in thecontext of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “about” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other. It should be appreciated that relativepositional terms such as “forward,” “aft,” “upper,” “lower,” “above,”“below,” and the like are with reference to normal operational attitudeand should not be considered otherwise limiting.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed:
 1. A core for use in casting an internal coolingcircuit within a gas turbine engine component, the core comprising: abase core portion; and an additive core portion additively manufacturedto the base core portion.
 2. The core as recited in claim 1, wherein thebase core portion is manufactured of a first material and the additivecore portion is manufactured of a second material, the first materialdifferent than the second material.
 3. The core as recited in claim 1,wherein the base core portion is manufactured of a ceramic material. 4.The core as recited in claim 1, wherein the base core portion is cast.5. The core as recited in claim 1, wherein the additive core portion isadditively manufactured of molybdenum.
 6. The core as recited in claim1, wherein the additive core portion forms a multiple of trailing edgecrossovers.
 7. The core as recited in claim 1, wherein the additive coreportion is additively manufactured onto the base core portion.
 8. Thecore as recited in claim 7, wherein the additive core portion includes amultiple of pin-shaped protrusions.
 9. The core as recited in claim 8,wherein each of the multiple of pin-shaped protrusions are additivelymanufactured into a respective recess formed in the base core portion.10. The core as recited in claim 1, wherein the base core portion ismore crushable than the additive core portion.
 11. A method ofmanufacturing a core for use in casting an internal cooling circuitwithin a gas turbine engine component, the method comprising: additivelymanufacturing an additive core portion to a base core portion.
 12. Themethod as recited in claim 11, further comprising casting the base coreportion.
 13. The method as recited in claim 11, further comprisingcasting the base core portion of a first material and additivelymanufacturing the additive core portion of a second material, the firstmaterial different than the second material.
 14. The method as recitedin claim 11, further comprising forming the additive core portion toform a multiple of trailing edge crossovers.
 15. The method as recitedin claim 11, further comprising additively manufacturing the additivecore portion to form a multiple of pin-shaped protrusions.
 16. Themethod as recited in claim 11, further comprising forming the additivecore portion onto the base core portion.
 17. The method as recited inclaim 11, further comprising firing the additive core portion and thebase core portion.
 18. The method as recited in claim 11, furthercomprising additively manufacturing the additive core portion to form atrailing edge portion of the core.
 19. The method as recited in claim11, further comprising positioning the core within a shell.
 20. Themethod as recited in claim 19, wherein the additive core portion atleast partially contacts the shell.